Manufacturing process for a high strength work hardened product made of AlZnMgCu alloy

ABSTRACT

The purpose of the invention is a process for the manufacture of a work-hardened product made of a high mechanical strength Al—Zn—Mg—Cu aluminium alloy consisting of:  
     casting an ingot made of an alloy with composition (% by weight) Zn=9.0-11.0, Mg=1.8-3.0; Cu=1.2-2.6 at least one of the elements Mn (0.05-0.4), Cr (0.05-0.3), Zr (0.05-0.20), Hf (0.05-0.5), V (0.05-0.3), Ti (0.01-0.2) and Sc (0.05-0.3), the remainder being made of aluminium and inevitable impurities,  
     possibly homogenisation of said ingot,  
     hot transformation of said ingot by rolling, extrusion or forging,  
     solution heat treatment and quenching of the product obtained,  
     possibly controlled stretching with a permanent set between 1 and 5%,  
     annealing of the product at a temperature and with a duration such that the product reaches the maximum compression yield stress in the L direction.  
     The invention is applicable particularly to upper wing members of aircrafts.  
                       Légende des figures                       Français   Anglais                   Seuil CSC   CSC threshold         Température (° C.)   Temperature (° C.)         Durée de revenu (h)   Annealing duration (h)         Mona-palier   Single step         Bi-palier   Two-step         Temps équivalent de revenu   Equivalent annealing time         à 120° C. (h)   at 120° C. (h)         Temps équivalent a 120° C.   Equivalent time at 120° C.         (h)   (h)         Tri-palier A   Three-step A         Tri-palier B   Three-step B         Temps équivalent de revenu   Equivalent annealing time         à 120° C. (h)   at 120° C. (h)

BACKGROUND OF THE INVENTION

[0001] 1. Field of the Invention

[0002] The invention relates to the manufacture of work-hardenedproducts by rolling, extrusion or forging made of a high mechanicalstrength aluminium alloy of the AlZnMgCu type, particularly productsused for aeronautical construction and particularly for upper wingmembers of aircrafts.

[0003] 2. Background Art

[0004] Al—Zn—Mg—Cu type alloys have been used for aircraft constructionfor more than 50 years, and particularly for upper wing members. Thus,7075, 7178, 7050, 7150 alloys and more recently 7055 and 7449 alloyshave been used. These alloys have usually been used in the T6 temper, inother words annealing corresponding to the maximum tensile yieldstrength, or an over annealed temper T76, T79 or T77 to obtain bettercorrosion resistance. As an illustration of this state of the art, theseare patents EP 0020505 by Boeing related to the 7150 alloy, U.S. Pat.Nos. 4,477,292, 4,832,758, 4,863,528 and 5,108,520 by Alcoa on the T77treatment, Alcoa' patent EP 0377779 dealing with a manufacturing processfor the 7055 alloy, and the applicant's patent application EP 0670377describing a process for the manufacture of plates made of 7449 alloy.

[0005] The properties of the 7449 alloy developed by the applicant forplates intended for use on upper wing members have been studied in thepaper written by T. Warner et al. “Aluminium alloy developments foraffordable airframe structures”, Conference on Synthesis, Processing andModelling of Advanced Materials, ASM International, Paris, Jun. 25-27,1977, pp 77-88. FIG. 2 in the article reproduced as FIG. 1 in thisapplication, shows typical properties of plates from 15 to 40 mm thickmade of this alloy, specifically the ultimate strength and the tensileyield strength in the L direction, the compression yield strength in theL direction and the stress corrosion limit (in the ST direction), in theT651 temper and in a T7×51 temper with improved corrosion resistance.This temper was identified in later publications by the same authors asT7951 (or T79511 for extruded products), for example in the paperwritten by F. Heymès et al. “New aluminium semi-products for airframeapplication”, METEC congress, Düsseldorf, June 1999 that uses the samefigure. FIG. 1 attached to this application shows that the compressionyield strength in the T79 temper is lower than the corresponding yieldstrength in the T6 temper. In the T7951 temper, plates made of 7449 havea 10% higher compression yield strength, better resistance toexfoliation corrosion and to stress corrosion and fatigue than platesmade of 7150 in the T651 temper that are usually used for upper wingmembers of commercial aircrafts, without any reduction in the toleranceto damage.

[0006] In summary, the state of the art shows firstly that themechanical compression strength is an essential property for upper wingmembers, and also that manufacturers of high strength alloys offerproducts for this application eight in the T6 temper corresponding tothe maximum tensile yield strength, or an over annealed T7 temper withbetter corrosion resistance but with lower mechanical strength.

BRIEF DESCRIPTION OF THE INVENTION

[0007] The purpose of the invention is to further improve the mechanicalcompression strength of products made of high strength 7000 alloysintended particularly for upper wing members of aircrafts without losingany other usage properties.

[0008] The purpose of the invention is a process for manufacturing awork-hardened product made of a high mechanical strength Al—Zn—Mg—Cualuminium alloy comprising:

[0009] casting an ingot made of an alloy with composition (% by weight)Zn (7.0-11.0), Mg (1.8-3.0), Cu (1.2-2.6), at least one of the elementsMn (0.05-0.4), Cr (0.05-0.30), Zr (0.05-0.20), Hf (0.05-0.5), V(0.05-0.3), Ti (0.01-0.2) and Sc (0.05-0.3), the remainder being made ofaluminium and inevitable impurities,

[0010] possibly homogenisation of said ingot,

[0011] hot transformation of said ingot by rolling, extrusion orforging,

[0012] solution heat treatment and quenching of the product obtained,

[0013] possibly controlled stretching with a permanent set between 1 and5%,

[0014] annealing of the product at a temperature and for a duration suchthat the product reaches the peak compression yield strength in the Ldirection.

[0015] Another purpose of the invention is the rolled, extruded orforged product obtained by said process.

[0016] Another purpose of the invention is the structural element formechanical construction, and particularly for aeronautical construction,made from at least one rolled, extruded or forged product obtainedaccording to said process.

BRIEF DESCRIPTION OF THE FIGURES

[0017]FIG. 1 shows typical properties, namely the tensile yield strength(L direction), the ultimate tensile strength (L direction), thecompression yield strength (L direction) and the stress corrosionthreshold (ST direction) for plates between 15 and 40 mm thick made ofthe 7150-T651, 7449-T651 and 7449-T7951 alloys according to prior art.

[0018]FIG. 2 shows the annealing time-temperature domain in the processaccording to the invention.

[0019]FIG. 3 shows the ultimate strength and the tensile yield strengthof 38 mm thick plates made of 7449 alloy in example 1 as a function ofequivalent annealing time at 120° C. for different annealingtemperatures.

[0020]FIG. 4 shows the ultimate tensile strength (L direction) and thetensile and compression yield strength (L direction) of 38 mm platesmade of 7449 alloy in example 2 as a function of the equivalentannealing time at 120° C.

[0021]FIG. 5 shows the compression yield strength of plates made ofalloys A and B in example 3 as a function of the equivalent annealingtime at 120° C.

DETAILED DESCRIPTION OF THE INVENTION

[0022] The invention is based on demonstrating a shift between the peakmechanical tensile strength obtained with annealing, corresponding towhat is usually called the T6 temper, and the peak mechanicalcompression strength. Although it has been known for a long time thatthe upper wing members are stressed mainly in compression, and thattherefore the compression yield strength controls the dimensions ofstructural elements of this part of the wing, metallurgists have alwaysused the tensile strength to define the T6 temper of maximum strengthachieved in annealing.

[0023] The inventors have found that there is a metallurgical statebetween tempers T6 and T79 according to prior art, in which thecompression yield strength passes through a peak of between 20 and 25MPa above the compression yield strength values of these two tempers.

[0024] The process according to the invention is applicable toAl—Zn—Mg—Cu type alloys with a high zinc content between 7 and 11%, witha magnesium content between 1.8 and 3%, and preferably between 1.8 and2.4%, and a copper content between 1.2 and 2.6% and preferably between1.6 and 2.2. The invention is not particularly useful for a zinc contentbelow 7% since these types of alloy are not longer used in aeronauticalconstruction for the fabrication of structural elements stressed incompression. For zinc contents higher than 11%, difficulties areencountered during industrial casting of rolling ingots or billets largeenough for the production of plates, sections or forged parts suitablefor the manufacture of the said structural elements.

[0025] The process according to the invention is applicable particularlyfor alloys used for manufacturing elements of the upper wing members ofaircrafts, for example 7055, 7349 and 7449 alloys in the form of workhardened products, in other words rolled, extruded or forged products.This process comprises the manufacture of an ingot, namely a rollingingot for rolled products, a billet or extrusion ingot for extrudedproducts or a forging ingot for forged products, in a known manner. Thisingot is preferably homogenised at a temperature close to the incipientmelting temperature of the alloy, as described in patent application EP0670377. It is then transformed by hot rolling, extrusion or forging, tothe required dimension. The product obtained is solution heat treatedalso at a temperature fairly close to the incipient melting temperatureof the alloy, this temperature being controlled by differentialenthalpic analysis. Solution heat treatment is followed by quenching,usually in cold water. The quenched product is preferably subjected tocontrolled stretching with a permanent set of between 1 and 5%.

[0026] The product is then annealed to obtain the peak compression yieldstrength in the L direction. Annealing may be done in a single step, inother words include a temperature rise gradient that may or may not belinear as a function of time, a period of time at a constant temperaturewithin the limit of the temperature tolerance of the furnace used, andcooling down to ambient temperature. For single step annealing, theconstant temperature is between 120and 150° C. with a duration withinthe parallelogram AEFG in FIG. 2 and preferably between 120 and 145° C.with a duration within the parallelogram ABCD in FIG. 2. The latterannealing process is a particularly preferred embodiment of thisinvention. For example, it can be used for products made from the 7449and 7439 alloys. Annealing may also be done in two steps, with a firststep at a temperature between 80 and 120° C., and a second step at ahigher temperature between 120 and 160° C. It may also be done in threesteps, with a first and a second step within the same limits as for thetwo-step annealing, and a third step at a lower temperature than thesecond step, between 100 and 140° C. Considering the time necessary fortemperature rises in industrial furnaces, it is difficult to envisagesteps lasting less than 2 h, and preferably less than 5 h.

[0027] In all cases, the two parameters (temperature and duration) canbe converted to a single parameter, the equivalent time at 120° C.defined by the following formula:${T({eq})} = \frac{\int{{\exp \left( {{- 16000}/T} \right)}{t}}}{\exp \left( {{- 16000}/T_{ref}} \right)}$

[0028] where T is the temperature of the annealing step in °K, t is thetreatment duration in hours and T_(ref) is the reference temperature inthis case assumed to be 120° C., namely 393°K. The equivalent annealingtime at 120° C. is between 100 and 250 h, and 50 to 200 h more than theequivalent annealing time for T651 annealing. The annealing timenecessary to reach the peak compression yield strength depends on thecomposition of the alloy and particularly the Cu/Mg ratio, the necessaryduration increasing with this ratio.

[0029] The work hardened product, and particularly the rolled, extrudedor forged product obtained using the process according to the invention,can advantageously be used for the manufacture of structural elements,particularly for aeronautical construction. Due to the increase in thecompression yield strength resulting from the process according to theinvention, a structural element manufactured from at least one extruded,rolled or forged product according to the invention has a betterresistance to compression loads than a structural element with the samedimensions made from work hardened, extruded or forged productsaccording to prior art. In one preferred embodiment of the invention,the structural element is an upper wing member of an aircraft.

EXAMPLES Example 1

[0030] 38 mm thick plates are made from 7449 alloy. The composition ofthe alloy is (% by weight) Zn=8.11, Mg=2.19, Cu=1.94, Si=0.04, Fe=0.07,Zr=0.09, Cr=0.06, Ti=0.025, the remainder aluminium and impurities(<0.05 each).

[0031] Plates have been pre-widened to increase the plate width from1100 mm to 2500 mm, hot rolled to 38 mm with an exit temperature of 378°C., solution heat treatment at 475° C., quenching with cold water andcontrolled stretching to 2.8% permanent set after waiting for 1 h afterquenching.

[0032] Samples taken from the mid-thickness of the plates were subjectedto 11 different single-step or two-step type annealings as shown intable 7. The rise and fall gradients between steps being 16° C./h and65° C./h respectively, corresponding to the rates observed forindustrial heat treatment furnaces. For each annealing, the equivalenttime at 120° C. t_(eq) is calculated using the following formula:${t({eq})} = \frac{\int{{\exp \left( {{- 16000}/T} \right)}{t}}}{\exp \left( {{- 16000}/T_{ref}} \right)}$

[0033] where T is the temperature of the annealing step in °K, t is thetreatment duration in hours and T_(ref) is the reference temperature, inthis case assumed to be 120° C., namely 393°K.

[0034] The 11 tested annealings were between annealing T651 andannealing T7951 according to prior art, and their parameters and thecorresponding equivalent times are shown in table 1.

[0035] In each case, the static tensile properties in the L directionwere measured (ultimate tensile strength R_(m), tensile yield strengthR_(0.2) and elongation A) on TOR 6 test pieces taken from the centralpart of the plates. The results are the average of at least twomeasurements and are shown in table 1 and in FIG. 3. TABLE 1 Teq atAnnealing Annealing 120° C. R_(0.2(L)) R_(m(L)) A 1^(st) step 2^(nd)step (h) (MPa) (MPa) (%) 24 h- 24 617 661 12 120° C. 48 h- 48 623 661 12120° C. 96 h- 96 624 655 12 120° C.  6 h-135° C. 29 616 655 12 12 h- 55619 651 11 135° C. 24 h- 108 619 651 11 135° C. 48 h- 215 611 642 11135° C. 24 h-  5 h-150° C. 125 620 649 11 120° C. 24 h-  9 h-150° C. 196613 642 11 120° C. 24 h- 13 h- 265 607 636 10 120° C. 150° C. 24 h- 17h- 336 595 627 10 120° C. 150° C.

[0036] It is found that close to the peak, the annealing processes withthe lowest temperature, in other words 120° C., give the highest valuesof R_(0.2) and R_(m). For two-step annealing processes, this effect iscontrolled by the temperature of the second step. Furthermore, the peaksfor R_(0.2) and R_(m) are similar, but not at exactly the same location.The T651 peak treatment can be defined as being the treatment thatresults in values of R_(0.2) and R_(m) within 5 MPa of the maximumpotential value, while remaining industrially acceptable. In this case,it is a 48 h treatment at 120° C.

Example 2

[0037] Samples taken from 38 mm thick plates made of 7449 alloy withcomposition of Zn=8.38, Mg=2.15, Cu=1.96, Si=0.04, Fe=0.06, Zr=0.11, theremainder aluminium and impurities (<0.5% each) are made in exactly thesame way as in example 1.

[0038] Eight different annealings were carried out on these samples,between annealing T651 defined in example 1 and annealing T7951. Thetemperatures and durations of these eight annealings and thecorresponding equivalent times at 120° C. are shown in table 2. TABLE 2Annealing Parameters Equivalent time A (T651) 48 h-120° C.  48 B 12h-135° C.  46 C 18 h-135° C.  78 D 24 h-135° C. 102 E 30 h-135° C. 130 F24 h-120° C. + 5.5 h-150° C. 130 G 24 h-120° C. + 11 h-150° C.  222 H(T7951) 24 h-120° C. + 17 h-150° C.  321

[0039] In addition to the mechanical tensile properties, the compressionyield strength was measured in the L direction on 13 mm diameter and 25mm long test pieces, and the electrical conductivity was measured onsamples taken from the surface. The average results of the twomeasurements are shown in table 3 and in FIG. 4, for R_(m) and R_(0.2)in tension, and R_(0.2) in compression. TABLE 3 R_(m) ten R_(0.2) ten AR_(0.2) comp Conduct Annealing (MPa) (MPa) (%) (MPa) (MS/m) A 633 67612.4 596 18.4 B 639 673 11.8 599 18.7 C 637 668 12.0 611 19.1 D 634 66311.0 614 19.7 E 633 663 10.5 615 20.0 F 635 662 11.2 613 20.1 G 619 64810.5 608 21.2 H 597 621 10.7 590 21.9

[0040] It is seen that the annealing that gives the peak compressionyield strength (L direction) is at an equivalent time of the order of150 h, in other words at an equivalent time intermediate between T651annealing and T7951 annealing. The useful range is between 100 and 250 hof equivalent time at 120° C., which is 50 to 200 h more than for a T651annealing. This annealing that results in the compression peak gives animprovement of 19 MPa compared with a T651 annealing and 25 MPa comparedwith a T7951 annealing.

Example 3

[0041] Plates made of two 7449 alloys with the thicknesses andcompositions indicated in table 4, were made in the same way as in theprevious examples as far as quenching. TABLE 4 e Alloy (mm) Si Fe Cu MgZn Zr Ti A 30 0.049 0.075 1.87 2.35 8.38 0.11 0.03 B 23 0.045 0.068 1.952.27 8.31 0.10

[0042] These plates were annealed as described in table 5, the first 11annealings being for alloy A and the last 7 for alloy B. The compressionyield strength R_(0.2) in the L direction, and the modulus of elasticityin compression also in the L direction, were measured on 13 mm diameterand 25 mm long test pieces taken from the central part of the plates.The results are shown in table 5, and the yield strength results areshown in FIG. 5 as a function of the equivalent annealing time at 120°C. TABLE 5 Annealing Annealing Annealing R0.2 comp Modulus 1^(st) step2^(nd) step 3^(rd) step (MPa) (MPa) 24 h-80° C. 24 h-135° C. 605 7028124 h- 24 h-135° C. 602 71200 100° C. 24 h- 24 h-135° C. 607 72335 120°C. 24 h- 18 h-140° C. 603 70598 100° C. 24 h-  7 h-150° C. 601 70618100° C. 24 h- 2.5 h- 607 72302 100° C. 160° C. 24 h- 30 h-140° C. 60072806 100° C. 24 h- 18 h-140° C. 24 h-120° C. 616 71621 100° C. 24 h-  7h-150° C. 24 h-120° C. 615 70862 100° C. 24 h- 2.5 h- 24 h-120° C. 62272569 100° C. 160° C. T7951 587 24 h-80° C. 24 h-135° C. 635 72910 24 h-24 h-135° C. 611 72222 120° C. 24 h- 18 h-140° C. 614 73244 100° C. 24h-  7 h-150° C. 610 72349 100° C. 24 h- 30 h-140° C. 596 70181 100° C.24 h-  7 h-150° C. 24 h-120° C. 621 71303 100° C. T7951 598

[0043] It is found that the peak compression yield strength occurs foran equivalent annealing time at 120° C. between 100 and 200 h, and thatthree-step annealing processes give higher values. Furthermore, it isfound that the compression yield strength is about 15 MPa better thanfor the T7951 annealing for two-step annealing processes, and about 25MPa better for three-step annealing processes.

What I claim is:
 1. Process for the manufacture of a work hardenedproduct made of a high mechanical strength Al—Zn—Mg—Cu aluminium alloycomprising: casting an ingot made of an alloy with composition (% byweight) Zn=7.0-11.0, Mg=1.8-3.0; Cu=1.2-2.6 at least one of the elementsMn (0.05-0.4), Cr (0.05-0.3), Zr (0.05-0.20), Hf (0.05-0.5), V(0.05-0.3), Ti (0.01-0.2) and Sc (0.05-0.3), the remainder being made ofaluminium and inevitable impurities, possibly homogenisation of saidingot, hot transformation of said ingot by rolling, extrusion orforging, solution heat treatment and quenching of the product obtained,possibly controlled stretching with a permanent set between 1 and 5%,annealing of the product at a temperature and with a duration such thatthe product reaches the maximum compression yield strength in the Ldirection.
 2. Process according to claim 1, wherein the magnesiumcontent of the alloy is between 1.8 and 2.4%.
 3. Process according toclaim 1, wherein the copper content of the alloy is between 1.6 and2.2%.
 4. Process according to claim 1, wherein the magnesium content ofthe alloy is between 1.8 and 2.4%, and the copper content is between 1.6and 2.2%.
 5. Process according to claim 1, wherein the alloy is 7349 or7449.
 6. Process according to claim 1, wherein the alloy is
 7055. 7.Process for the manufacture of a work hardened product made of a highmechanical strength Al—Zn—Mg—Cu aluminium alloy comprising: casting aningot made of an alloy with composition (% by weight) Zn=7.0-11.0,Mg=1.8-3.0, Cu=1.2-2.6 at least one of the elements Mn (0.05-0.4), Cr(0.05-0.3), Zr (0.05-0.20), Hf (0.05-0.5), V (0.05-0.3), Ti (0.01-0.2)and Sc (0.05-0.3), the remainder being made of aluminium and inevitableimpurities, possibly homogenisation of said ingot, hot transformation ofsaid ingot by rolling, extrusion or forging, dissolution and quenchingof the resulting product, possibly controlled stretching with apermanent set between 1 and 5%, single step annealing at a temperatureand with a duration included within the parallelogram AEFG, in which thevertices in the temperature-duration diagram have the followingcoordinates: A: 120° C.-100 h E: 145° C.-5 h F: 150° C.-40 h G: 120°C.-700 h.
 8. Process according to claim 7, wherein the annealing is asingle step annealing at a temperature and with a duration within theparallelogram ABCD, in which the vertices in the temperature-durationdiagram have the following coordinates: A: 120° C.-100 h B: 145° C.-9 hC: 145° C.-22 h D: 120° C.-230 h.
 9. Process according to claim 1,wherein the equivalent annealing time at 120° C. is between 100 and 250h.
 10. Process according to claim 1, wherein the equivalent annealingtime at 120° C. is 50 to 200 h longer than the time corresponding totemper T651.
 11. Process according to claim 1, wherein said annealing isa two-step annealing comprising a first step at a temperature between80° C. and 120° C., and a second step at a temperature between 120° C.and 160° C., and wherein said equivalent annealing time at 120° C. isbetween 100 and 250 h.
 12. Process according to claim 1, wherein saidannealing is a three-step annealing comprising a first step at atemperature between 80° C. and 120° C., a second step at a temperaturebetween 120° C. and 160° C., and a third step at a lower temperaturethan the second step and between 100° C. and 140° C., and wherein theequivalent annealing time at 120° C. is between 100 and 250 h. 13.Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim 1.14. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim 2.15. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim 3.16. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim 4.17. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim 5.18. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim 6.19. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim 7.20. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim 8.21. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim 9.22. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim10.
 23. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim11.
 24. Structural element for mechanical construction and particularlyaeronautical construction, manufactured from at least one rolled,extruded or forged product obtained by the process according to claim12.